This invention relates generally to a main combustor in a gas turbine engine, and, more particularly, to a segmented, zoned fuel injection system for use within the combustor of the gas turbine engine.
Fuel is generally introduced into the combustion chamber of a combustor in a gas turbine engine by means of nozzles. Unfortunately, conventional nozzle fuel injectors generally have no accommodation for radial variation of fuel flow, since the conventional design is made up of a single, point source for the fuel nozzle. In addition, the conventional fuel nozzles are located two to three inches apart circumferentially.
Although such an arrangement can provide a desired flow variation circumferentially by utilizing a high or low flowing nozzle; such nozzles flow high or low during all flight conditions and flow requirements. For example, high local fuel flow at one flight point such as at ignition, would be a drawback at other flight points such as at maximum power where high local fuel flow creates a high pattern factor. More specifically, fuel spray discharge or injection systems of the past left much to be desired if a specific quantity of fuel were needed to be introduced at a specific location in the combustion chamber, if local circumferentially high temperatures at the turbine inlet were to be eliminated or if adjustment of fuel flow in the area of the igniter to aid ignition was desired.
It is therefore clearly evident from the above recited problems that it would be highly desirable to provide a fuel spray injection system which is capable of effective and efficient operation within a main combustor of a gas turbine engine.